FIG. 1 shows a gas turbine engine 10. The engine 10 comprises an air intake 12 and a propulsive fan 14 that generates two airflows A and B. The gas turbine engine 10 comprises, in axial flow A, an intermediate pressure compressor 16, a high pressure compressor 18, a combustor 20, a high pressure turbine 22, an intermediate pressure turbine 24, a low pressure turbine 26 and an exhaust nozzle 28. A nacelle 30 surrounds the gas turbine engine 10 and defines, in axial flow B, a bypass duct 32.
Some of the components of the gas turbine engine may comprise nickel alloys, such as nickel super-alloys. One example nickel alloy component comprises a turbine blade 34 shown in FIG. 2. The blade 34 forms part of the high pressure turbine 22 and comprises a suction side 36 and a pressure side 38, as well as leading and trailing edges 40, 42, a root 44 at a radially inner end, and a tip 46 at a radially outer end.
The turbine blade 34 includes a plurality of cooling holes 48 leading from an internal passage 50, which extends from the root 44 to the tip 46. The holes 48 provide cooling air provided from one of the compressors 16, 18 to prevent softening of the metal in use, since gases flowing over the turbine blades 34 are generally at a very high temperature.
The positioning and size of the cooling holes 48 must be selected to ensure that no part of the blade 34 is subject to excessively high temperatures during use, while minimising use of cooling air, since the cooling air bled from the compressors 16, 18 is essentially lost to the thermodynamic cycle of the engine 10. Essentially therefore, excessive use of cooling air leads to increased specific fuel consumption. During the design stage of a new blade 34, it is therefore generally necessary to test a blade 34 to determine the amount of cooling provided by the cooling holes 48 during different conditions.
In one known testing method, steam is forced through the passage 50 to emerge from the cooling holes 48, and the flow through the holes 48 is measured. The holes 48 may then be increased in size by a small amount using a mechanical, electrochemical or acid etching process, and the blade 34 is then tested again. Other surface features, such as the external aerodynamic surfaces 36, 38, 40, 42 may also require adjustment by removal of small amounts of material following testing.
However, prior methods of removing small amounts of material from the blade 34 to enlarge the holes 48 or change the dimensions of other surface features produce unwanted disruption of the surface properties of the blade, which therefore decreases the accuracy of the tests. It is also difficult using conventional mechanical methods to remove sufficiently precise small amounts of material. Electrochemical processes are generally energy intensive and slow. Acid etching requires the use of hazardous chemicals, and thereby introduces hazards to the user, and increased costs due to the requirements for safety equipment and safe disposal of chemicals.
The present invention describes a method of machining a surface feature of a nickel alloy component of a gas turbine engine, which seeks to overcome some or all of the above problems.